Active force generation system for minimizing vibration in a rotating system

ABSTRACT

A method and device for reducing vibratory noise in a system with an integral rotating member includes independently operable drive systems for controlling the angular velocity of at least two independently rotatable masses. Control signals manipulate the drive system to rotate each mass at optimal speed, direction and phase to reduce noise induced in the system by the rotating member.

REFERENCE TO RELATED APPLICATIONS

The present invention is a Continuation of Divisional Application No.12/352676, filed Jan. 13, 2009, which is a Divisional Application ofU.S. patent application Ser. No. 10/685215, filed Oct. 14, 2003.

BACKGROUND

This invention relates to vibration isolators, and more particularly, toan isolation system for minimizing in-plane vibrations produced in arotating system of a rotary-wing aircraft, and still more particularly,to an isolation system that minimizes system weight, aerodynamic drag,and complexity while concomitantly providing active control andadjustment during operation for optimal efficacy across a wide spectrumof operating speeds.

Vibration isolation or absorption is oftentimes desirable for nulling orcanceling vibrations associated with a rotating system. Such vibrations,when left unattenuated or unabated, may lead to structural fatigue andpremature failure of system components. Furthermore, inasmuch as suchvibrations may be transmitted through adjacent support structure to, forexample, an aircraft avionics bay, areas occupied by passengers, orother components and cabin area remote from the source of the vibrationwhich may also be subject to these same potentially damaging ordisturbing vibrations (albeit perhaps lower in amplitude due to energyabsorption by the interconnecting structure). Consequently, it is mostdesirable to isolate or absorb these vibrations at or near the source ofthe vibration in the rotating system.

One application which best exemplifies the need for and advantagesderived from vibration isolation/absorption devices is the main torquedriving hub of a helicopter rotor system. Typically, the main rotor of ahelicopter, which comprises a central torque drive hub member fordriving a plurality of lift producing rotor blades, is subject to avariety of aerodynamic and gyroscopic loads. For example, as each rotorblade advances or retreats relative to the freestream airflow, itexperiences a sharp rise and fall of in-plane aerodynamic drag.Furthermore, as the tip of each rotor blade advances with eachrevolution of the rotor system, the relative velocity of the blade tipapproaches supersonic Mach numbers. As such, large variations occur inthe various coefficients which define blade performance (e.g., moment,lift and drag coefficients). Moreover, gyroscopic and Coriolus forcesare generated causing the blades to “lead” or “lag” depending uponcyclic control inputs to the rotor system. All of the above generatesubstantial in-plane and out-of-plane vibrations which, if notsuppressed, isolated or otherwise abated, are transmitted to the cockpitand cabin, typically through the mounting feet of the helicopter mainrotor gearbox.

Various vibration isolation systems have been devised tocounteract/oppose and minimize these in-plane and out-of-planevibrations. Mast-mounted vibration isolators suppress or isolatein-plane vibrations at a location proximal to the source of suchin-plane vibrations whereas transmission, cabin or cockpit absorbersdampen or absorb out-of-plane vibrations at a location remotely disposedfrom the source. Inasmuch as the present invention relates to theisolation of in-plane vibrations, only devices designed tocounteract/oppose such vibrations will be discussed herein.

Some mast-mounted vibration isolators have a plurality of resilient arms(i.e., springs) extending in a spaced-apart spiral pattern between a hubattachment fitting and a ring-shaped inertial mass. Several pairs ofspiral springs (i.e., four upper and four lower springs) are mounted toand equiangularly arranged with respect to both the hub attachmentfitting and the inertial mass so as to produce substantially symmetricspring stiffness in an in-plane direction. The spring-mass system, i.e.,spiral springs in combination with the ring-shaped mass, is tuned in thenon-rotating system to a frequency equal to N * rotor RPM (e.g., 4P fora four-bladed rotor) at normal operating speed, so that in the rotatingsystem it will respond to both N+1 and N−1 frequency vibrations (i.e.,3P and 5P for a four-bladed rotor). N is the number of rotor blades.

While these spiral spring arrangements produce a relatively small widthdimension (i.e., the spiraling of the springs increases the effectivespring rate), the height dimension of each vibration isolator isincreased to react out-of-plane loads via the upper and lower pairs ofspiral springs. This increased profile dimension increases the profilearea, and consequently the profile drag produced by the isolator. Thespiral springs must be manufactured to precise tolerances to obtain therelatively exact spring rates necessary for efficient operation suchthat manufacturing costs may be increased. Furthermore, these vibrationisolators are passive devices which are tuned to a predeterminedin-plane frequency. That is, the vibration isolators cannot be adjustedin-flight or during operation to isolate in-plane loads which may varyin frequency depending upon the specific operating regime.

Another general configuration of isolator known as a “bifilar” aremast-mounted vibration isolators having a hub attachment fittingconnected to and driven by the helicopter rotorshaft, a plurality ofradial arms projecting outwardly from the fitting and a mass coupled tothe end of each arm via a rolling pin arrangement. That is, a pin rollswithin a cycloidally shaped bushing thereby permitting edgewise motionof each mass relative to its respective arm. The geometry of the pinarrangement in combination with the centrifugal forces acting on themass (imposed by rotation of the bifilar) results in an edgewiseanti-vibration force at a 4 per revolution frequency which isout-of-phase with the large 4 per revolution (or “4P” as it is commonlyreferred to as helicopter art) in-plane vibrations of the rotor hub fora 4 bladed helicopter. The frequency of 4P is the frequency as observedin a non-rotating reference system.

More specifically, pairs of opposed masses act in unison to produceforces which counteract forces active on the rotor hub. In FIG. 1, aschematic of a pair of bifilar masses, at one instant in time, aredepicted to illustrate the physics of the device. Therein, the massesMI, MII are disposed at their extreme edgewise position within each ofthe respective cycloidal bushings BI, BII. The masses MI, MII producemaximum force vectors F/2, which produce a resultant vector F at thecenter, and coincident with the rotational axis, of the rotating system.The combined or resultant force vector F is equal and opposite to themaximum vibratory load vector P active on the rotor at the same instantof time. This condition, when the bifilar produces an equal and oppositeforce F that opposes the rotor load P, reflects ideal operation of thebifilar. Excessive bifilar damping or manufacturing imperfections willcause the bifilar output force F to differ from the disturbing force Pproduced by the rotor either in magnitude or phase best suited tonullify the rotor loads. This condition may cause unwanted fuselagevibration. It will also be appreciated that for the masses to producethe necessary shear forces to react the in-plane vibratory loads of therotor system, counteracting bending moments are also produced. Theseforce couples impose large edgewise bending loads in the radial arms,and, consequently, the geometry thereof must produce the necessarystiffness (EI) at the root end of the arms. As such, these increasedstiffness requirements require the relatively large and heavy bifilararms.

While the bifilar system has proven effective and reliable, the weightof the system, nearly 210 lbs, is detrimental to the overall liftingcapacity of the helicopter. To appreciate the significance of theincreased weight, it has been estimated that for each pound ofadditional weight, direct operating cost of the helicopter may increaseby approximately $10,000.

Furthermore, the pin mount for coupling each mass to its respectiveradial arm routinely and regularly wear, thus requiring frequent removaland replacement of the cyclical bushings. This increases the DirectMaintenance Costs (DMC) for operating the helicopter, which contributes,to the fiscal burdens of the bifilar system and the helicopter.

Therefore, a need exists for an isolation system to reduce vibrations ina rotating system that isolates a wide spectrum of vibratory loads;especially large amplitude loads, minimizes system weight, reducesaerodynamic drag and reduces DMC.

SUMMARY

The present invention provides a vibration isolation system which iscontrollable for varying the range of isolation frequencies whichabsorbs large amplitude vibrations while minimizing system weight.

The vibration isolation system employs readily manufactured componentswhich is insensitive to damping and manufacturing imperfections.

BRIEF DESCRIPTION OF THE DRAWINGS

A more complete understanding of the present invention and the attendantfeatures and advantages thereof may be had by reference to the followingdetailed description when considered in conjunction with theaccompanying drawings wherein:

FIG. 1 is a schematic of a prior art bifilar isolation device forillustrating certain physical characteristics thereof.

FIG. 2 is a side sectional view of a helicopter main rotor, including amain rotor shaft having an isolation system according to the presentinvention mounted to the upper mast or shaft extension member of therotor.

FIGS. 3 a-3 c depict schematic views of various operating conditions ofthe inventive isolation system.

FIG. 4 is a side sectional view of one embodiment of the isolationdevice.

DETAILED DESCRIPTION

The isolation system of the present invention is described in thecontext of a helicopter rotor system, such as that employed in an ArmyBLACK HAWK helicopter produced by Sikorsky Aircraft Corporation. Oneskilled in the art, however, will appreciate that the present inventionhas utility in any rotating system which produces vibratory loads(noise). The invention is especially useful in rotating systems thatproduce large vibratory loads that vary depending upon differentoperating regimes or variable operating speeds.

Referring to FIG. 2, the vibration isolation system 10 is disposed incombination with a rotary-wing aircraft main rotor system 2 having amain rotor shaft 4 (rotating system member) that is driven about arotational axis 6 by a torque driving transmission 8. In the describedembodiment, the rotor system 2 includes a hub 12 having four radial armsthat mount to and drive each rotor blade 16. The vibration isolationsystem 10 is mounted to a flanged end 13 of the main rotor shaft 4through a hub attachment fitting 18. Vibratory forces active on the mainrotor system 2 are generated by a variety of factors, although thedominant vibrations originate from aerodynamic and/or gyroscopic forcesgenerated by each rotor blade 16. A four bladed rotor system produces 3Pvibratory loads, i.e., in a single revolution, the magnitude of the loadvector varies from a minimum to a maximum value three times in therotating frame of reference. This resolves into 4P vibration in thenon-rotating frame of reference due to the addition of the 1P rotorrotational speed. While a variety of factors influence the vibratoryspectrum of a rotor system, such 4P vibrations are generally a result ofeach rotor blade experiencing maximum lift when advancing and minimumlift when retreating.

Referring to FIGS. 2 and 4, the vibration isolation system 10 includestwo, essentially coplanar, masses M1, M2, a drive system 30 for drivingthe masses M1, M2 about the rotational axis 6 of the main rotor shaft 4,a control system 40 for issuing control signals to the drive system 30to control the rotational speed and relative angular position of themasses M1, M2 and a power source 50 for energizing the drive system 30and control system 40.

The masses M1, M2 are (i) disposed at a predetermined distance R fromthe main rotor shaft axis 6; (ii) driven in the same or opposingrotational direction as the main rotor shaft axis 6; and (iii) driven ata rotational speed at least 3P greater than the rotational speed 1P ofthe rotor shaft 4. In one embodiment, the drive system 30 includes apair of electric motors 34 a, 34 b for driving each of the masses M1, M2through a relatively small diameter, constant cross-section radial arm36 (shown schematically in FIG. 3 a-3 c). Moreover, the electric motors34 a, 34 b are independent of each other, e.g., may be driven atdifferent rotational speeds to enable variation of the isolation forcemagnitude and phase.

As shown in FIG. 4, the control system 40 requires a speed sensor 42 forissuing signals 42 s indicative of the rotational speed 1P of the rotorshaft 4, and a signal processing and amplifier 44, responsive to thespeed signals 42 s, to issue control signals 44 s to the drive system 30indicative of the rotational velocity and relative angular position ofeach of the masses M1, M2. While the speed sensor 42 may be a dedicatedunit for sensing rotor speed, the same information may be obtained froma transmission alternator or generator 50 which turns at a predefinedspeed multiple relative to the rotor speed. The alternator or generator50 supplies power to the controller-amplifier 44 through the slip ring54. Hence, the control system 40 may use voltage phase information fromsuch devices to issue the appropriate control signals to the drivesystem 30.

While the isolation system 10 may employ a control system 40 having apredefined schedule or model of the vibrations, e.g., at prescribedrotor speeds, another embodiment may also employ a vibration sensingdevice or system. As such, the control system 40 includes one or morevibration feedback sensors 51 for issuing vibration signals 51 sindicative of the vibrations (e.g., amplitude, frequency and phase) ofthe helicopter rotor hub 12. The control system 40, therefore, samplesvibration levels at predefined intervals or rates to identify atrend-positive (lower vibration levels) or negative (larger vibrationlevels). Accordingly, as vibration levels change, the control system 40issues modified signals 44 s to the drive system 30 until an optimumcombination of rotational speed, force magnitude and phase are achieved.

The isolation system 10 may be powered by any of a variety of knownmethods, especially methods which may require transmission from astationary to a rotating reference field. In the described embodimentshown in FIG. 4 the drive system 30 and control system 40, respectively,are powered by a 15 kVa generator 50 which provides a 115 volt potentialat 400 Hz and with 3 phases (typical AC power for helicopters). Power istransferred from the stationary system to the rotating system via aconventional cylindrical slip ring 54. Only a small amount of additionalweight is required inasmuch as the slip ring 54 is pre-existing and usedfor powering other systems e.g., rotor blade de-ice system. This slipring may also be used to communicate the control signals 42 s to thedrive system 30 when the control system 40 is mounted in the fuselagerather than on the rotor system 2.

In operation, the masses M1, M2 (shown in FIGS. 3 a-3 c) are driven bythe drive system 30 at a rotational speed greater than the rotationalspeed of the rotating system and appropriately positioned to yield aload vector P10 which is equal and opposite to the load vector P2produced by the rotor system 2. This counteracting load vector P10 canbe viewed as a vector which attempts to cancel or null the displacementof the rotor shaft 4. In the described embodiment, the masses turn at arotational speed.

Inasmuch as the drive system 30 is mounted directly to the rotatingshaft 4 of the rotor system 2, the drive system 30 need only drive themasses M1, M2 three additional revolution per cycle (for each revolutionof the rotor system) to achieve the desired 4P frequency. That is, sincethe masses M1, M2 are, in a rotating reference system, driven at onerevolution per cycle by the rotor system 2 itself, the drive means 30need only augment the rotational speed by the difference (4P−1P) toachieve the necessary 4P in the stationary reference system.

FIGS. 3 a-3 c depict various operating positions of the masses M1, M2 toemphasize the function and versatility of the isolation system 10.

In FIG. 3 a, the masses M1, M2 are essentially coincident and act inunison to produce a maximum force vector P10MAX.

In FIG. 3 b, the masses M1, M2 define a right angle (90 degrees)therebetween thereby producing a force vector P10MAX/(sqrt (2)) that isa fraction of the magnitude of the maximum force vector.

In FIG. 3 c, the masses M1, M2 define a straight angle (180 degrees) andare essentially opposing to cancel the vectors produced by each of themasses M1, M2 independently or individually.

In FIG. 4, the controller 40 issues signals to the drive system 30 to(a) drive the masses M1, M2 at a rotational speed greater than that ofthe rotating system and (b) produce a counteracting load of the correctmagnitude and phase to efficiently isolate vibrations.

The ability to independently vary the relative angular position of themasses M1, M2 is especially valuable in applications wherein themagnitude of the vibratory load active in/on the rotating system variesas a function of operating regime or operating speed. In a rotary-wingaircraft, for example, it is common to require the highest levels ofvibration isolation in high speed forward flight i.e., where the rotorblades are experiencing the largest differential in aerodynamic loadingfrom advancing to retreating sides of the rotor system. Consequently, itmay be expected that the drive system 30 produce the maximum load vectorP10MAX such as illustrated in FIG. 3 a. In yet another example, it isanticipated that the lowest levels of vibration isolation would occur ina loiter or hovering operating mode, where the rotor blades are exposedto the generally equivalent aerodynamic and gyroscopic affects.Consequently, it may be expected that the drive means 30 produce no or aminimum load vector P10MIN such as illustrated in FIG. 3 c.

Thus far, the discussion herein has concentrated on the rotational speedand angular position of the masses M1, M2 to produce vibrationisolation. While this feature of the invention is a primary aspect ofthe invention, the configuration of the inventive isolation system 10produces counteracting load vectors P10 which act though the rotationalaxis of the rotor shaft 4. That is, the line of action of the loadvector P10, whether the masses M1, M2 are coincident or opposing,intersects the rotational axis and produces pure radial loads. As such,the radial arms of the isolation system 10 are principally loaded intension rather than a combination of tensile and bending moment loads. Aconsequence of this loading condition is a reduction in system weightinasmuch as the radial arms 36 need not produce high edgewise strengthto react bending moment loads.

Furthermore, tensile loading in the radial arms 36 enables the use of aconstant-cross-section structure to react the centrifugal loads producedby each of the masses M1, M2. Moreover, directional strength materials(non-isotropic) may be employed such as unidirectional fiber reinforcedcomposites. As a result, the isolation device may be produced usingrelatively low cost manufacturing techniques and materials. For example,cylindrical raw material stock, cut to the proper length, may beemployed without secondary processing. Also, the use of unidirectionalcomposites enables yet further weight reduction.

Although the invention has been shown and described herein with respectto a certain detailed embodiment of a mast-mounted helicopter isolator,it will be understood by those skilled in the art that a variety ofmodifications and variations are possible in light of the aboveteachings. It is therefore to be understood that, within the scope ofthe appended claims, the present invention may be practiced otherwisethan as specifically described hereinabove.

1. A rotary-wing aircraft rotor system which rotates about an axis ofrotation, comprising: a rotor system having an N number of blades whichrotates about an axis of rotation at a rotational speed of 1P such thatsaid main rotor system produces NP vibrations; a sensor system whichsenses the NP vibrations; a multiple of masses coaxially disposed withsaid rotor system, said multiple of masses includes a first mass and asecond mass; a first drive system to spin said first mass about saidaxis of rotation at a first angular velocity; a second drive system tospin said second mass about said axis of rotation at a second angularvelocity; and a control system in communication with said sensor systemand said drive system, said control system operable to identifyvariations of the NP vibrations to control the angular velocity of atleast one of said multiple of masses to reduce the NP in-planevibrations.
 2. The system as recited in claim 1, wherein each of themultiple of masses are within a separate plane.
 3. The system as recitedin claim 1, wherein said drive system rotates at least one of saidmultiple of masses in a direction opposite to the rotational directionof said rotor system.
 4. The system as recited in claim 1, wherein saidmultiple of masses spin continuously while said multiple of massesreduce the NP vibrations.
 5. The system as recited in claim 1, whereinsaid control system utilizes a phase angle from a power source driven bysaid main rotor system as a phase angle reference to said controlsystem.
 6. The system as recited in claim 5, wherein said sensor systemis interconnected to said main rotor system to provide feedback signalsto said control system.
 7. The system as recited in claim 1, whereineach of said multiple of masses includes an eccentric mass definedrelative to said axis of rotation.
 8. A rotary-wing aircraft rotorsystem which rotates about an axis of rotation, comprising: a rotorsystem having an N number of blades which rotates about an axis ofrotation at a rotational speed of 1P, such that said main rotor systemproduces NP vibrations; a sensor system which senses the NP vibrations;a multiple of masses coaxially disposed with said rotor system; a drivesystem to independently spin each of said multiple of masses about saidaxis of rotation at an independent angular velocity; and a controlsystem in communication with said sensor system and said drive system,said control system operable to identify variations of the NP vibrationsto control the angular velocity of at least one of said multiple ofmasses to reduce the NP vibrations.
 9. The system as recited in claim 8,wherein said rotor system includes a rotary wing aircraft main rotorsystem.
 10. The system as recited in claim 8, wherein said drive systemrotates at least one of said multiple of masses in a direction oppositeto the rotational direction of said rotor system.
 11. The system asrecited in claim 8, wherein said drive system rotates at least one ofsaid multiple of masses at an angular velocity greater than an angularvelocity of said rotor system.
 12. The system as recited in claim 8,wherein said control system utilizes a phase angle from a power sourcedriven by said main rotor system as a phase angle reference to saidcontrol system.
 13. The system as recited in claim 12, wherein saidcontrol system communicates with a sensor system interconnected to saidmain rotor system to provide feedback signals of said NP vibrations tosaid control system.
 14. A vibration isolation system for reducingvibrations in a rotating system rotatable about an axis of rotation,comprising: a multiple of masses coaxially disposed about an axis ofrotation of a rotating system, each of said multiple of masses radiallyoffset from said axis of rotation; a drive system to independently spineach of said multiple of masses about said axis of rotation at anangular velocity; and a control system in communication with said drivesystem to control the angular velocity of each of said multiple ofmasses to reduce vibrations generated by the rotating system.
 15. Thesystem as recited in claim 14, wherein said rotor system includes arotary wing aircraft main rotor system.
 16. The system as recited inclaim 14, wherein said drive system spins at least one of said multipleof masses in a direction opposite to the direction of rotation of saidrotating system.
 17. The system as recited in claim 14, wherein saiddrive system spins at least one of said multiple of masses at an angularvelocity greater than an angular velocity of said rotating system. 18.The system as recited in claim 14, wherein said control system utilizesa phase angle from a power source as a phase angle reference to saidcontrol system.
 19. The system as recited in claim 14, wherein each ofsaid multiple of masses are mounted to an end of a radial arm.
 20. Thesystem as recited in claim 19, wherein said drive system includes amultiple of electric motors, each of said multiple of electric motorsspin one of said multiple of masses through said radial arm.
 21. Thesystem as recited in claim 20, wherein each of said multiple of electricmotors are independently operated to independently spin each of saidmultiple of masses at an independent angular velocity.
 22. The system asrecited in claim 14, wherein each of said multiple of masses define aneccentric mass relative to said axis of rotation.
 23. The system asrecited in claim 14, wherein said multiple of masses are located only onone side of said rotating system.
 24. The system as recited in claim 14,wherein said drive system spins each of said multiple of masses aboutsaid axis of rotation at an independent angular velocity.